In gas turbine engines, air is compressed at an initial stage, then is heated in combustion chambers, and the hot gas so produced passes to a turbine that, driven by the hot gas, does work which may include rotating the air compressor.
In a typical industrial gas turbine engine a number of combustion chambers combust fuel and hot gas flowing from these combustion chambers is passed via respective transitions (also referred to by some in the field as ducts and tail tubes) to respective entrances of the turbine. More specifically, a plurality of combustion chambers commonly are arranged radially about a longitudinal axis of the gas turbine engine, and likewise radially arranged transitions comprise outlet ends that converge to form an annular inflow of hot gas to the turbine entrance. Each transition exit is joined by a number of seals each of which bridges a gap between a portion of the exit and one or more turbine components. The latter, in various designs, are identified as row 1 vane segments. Adjacent component growth variances due to thermal expansion, mechanical loads, and vibrational forces from combustion dynamics all affect design criteria and performance of such a seal, referred to herein as a transition-to-turbine seal. Maintenance of component temperatures below particular limits is also desired and this may affect design of the seal and adjacent components. Consequently, the design of such seals has presented a challenge resulting in various approaches that attempt to find a suitable balance between seal cost, reliability, durability, installation and repair ease, performance, and effect on adjacent components.
For example, U.S. Pat. No. 6,751,962, issued Jun. 22, 2004 to Kuwabara et al., provides inclined cooling fluid holes drilled in a tail tube seal in addition to conventionally existing cooling fluid holes. These cooling fluid holes exit into the hot gas path, and are stated to cool the hot gas side of a downstream groove of the seal due to film effect. This is stated to increase reliability and decrease wear. A different approach is taken to cool the transition side of the seal in U.S. Pat. No. 6,769,257, issued Aug. 3, 2004 to Kondo et al. In this patent are disclosed cooling medium and heating medium channels provided in the outlet structure of a transition. Various embodiments are described that are stated to reduce the temperature difference of a flange formed at the downstream end of the transition, which attaches to a sealing component connecting to the turbine. Finally, in U.S. Pat. No. 6,860,108, issued Mar. 1, 2005 to Soechting et al., a seal was directed to prevent the outer and inner shrouds of the turbine's first stationary blade (i.e., row 1 vane segment) from heat damage and wear. The seal comprised a downstream portion having an inclined surface (inclining outwardly from the hot gas path) so that the cross-sectional area defined within the seal increased from an upstream point to a downstream point. Also, outlets for ejecting cooling air were provided that were disposed to release cooling air at the downstream end of the seal. Further, bleed holes were provided toward an upstream end section of the seal, near a front corner of the seal in the hot gas path. The latter are stated to “cool the film” [sic] of the parallel (non-inclined, more upstream) and the inclined (more downstream) surfaces of the seal that are in the hot gas path.
Despite the respective features of these and other transition-to-turbine seals and temperature equilibrating approaches known in the art, there remains a need for an improved transition-to-turbine seal.